A. Field of Invention
The present invention relates to a diffuser by which an airflow can be reduced from a higher velocity to a lower velocity, and more particularly to such a diffuser used in a wind tunnel to bring the airflow from a higher velocity, lower static pressure condition to a lower velocity, higher static pressure condition for recirculation through the wind tunnel.
B. Background Art
In a conventional wind tunnel where air is driven by a main drive fan through a test section of the tunnel, it is necessary to pass the air through at least two diffuser sections. There is a first diffuser section which receives relatively high velocity air from the test section to cause the air to decrease in velocity (with a consequent increase in static pressure) so that this air can more effectively be moved by the drive fan in its recirculating pattern in the wind tunnel. There is also a fan diffuser section which receives the air from the main drive fan and passes this air through a settling chamber from which the air passes through a constricted passageway section into the test section at a relatively higher velocity. In the settling chamber area, the air may be passed through a cooling device to extract from the air the heat that results from being recirculated through the wind tunnel. There may also be a flow straightening device which is often in the form of a plurality of elongated passageways to provide a more uniform flow into the test section.
In a conventional wind tunnel, the requirements of the diffuser sections dictate to a large extent the overall tunnel circuit length. The main function of the diffuser sections is to decelerate the flow from its maximum velocity in the tunnel section to much lower velocities, and to recover as much as possible of the flow's original kinetic energy in the form of increased pressure. In a conventional diffuser, the deceleration and the associated pressure increase must be gradual, lest the boundary area separate from the diffuser wall, causing large energy losses and flow unsteadiness. Conventional conical diffusers without active boundary layer control are thus limited to included angles (i.e. the total angle between the walls) between five and seven degrees and must therefor be quite long. While maintaining attached flow through good design practice avoids the very large losses that would occur if the flow separated, it does not eliminate energy losses. The first diffuser section downstream of the tunnel test section in particular typically accounts for a substantial fraction of the power which must be supplied by the tunnel drive system.
Active boundary layer control offers the prospect of dramatically reducing the length of the diffuser sections and possibly reducing the diffuser energy losses by half or more. These advantages would be of most value in a large, pressurized wind tunnel, and would thus be most likely to justify increased mechanical complexity. In such a facility, the walls of the tunnel circuit constitute a very expensive pressure vessel structure, and shortening the circuit thus could have a large impact on initial costs.
Also, in large tunnels the cost of energy is an important factor in operating costs and the reduction in energy losses could produce substantial savings.
The two forms of active boundary layer control that have been proposed for use in diffusers are suction and tangential blowing.
In boundary layer suction schemes some of the lowest energy air adjacent to the wall (i.e at the "bottom" of the boundary layer) is removed through perforations or slots in the wall, thus delaying the flow reversal that characterizes separation. The boundary layer is then able to withstand a more abrupt pressure rise, and the diffuser can be made shorter. Substantial shortening of the diffuser, however, requires that suction be applied to a substantial part of the surface over which the pressure gradient is felt, and therefore the suction must be distributed over multiple holes or slots. Complex ducting schemes are required to handle the suction air efficiently (i.e. to minimize further energy losses), increasing the complexity of the diffuser wall structure. Even with good duct design the suction air tends to have low energy, partly because only the lowest energy air in the boundary layer is removed, and partly because the small scale slots or holes which are used in the schemes have significant losses of their own. The pumping power to return the suction air to the tunnel circuit can therefore be relatively large.
In tangential blowing schemes, a high velocity jet of air (called a wall-jet) is introduced through a slot upstream of the region of pressure rise. This high energy air mixes with the boundary layer and enables it to overcome a much more abrupt pressure rise without separating. A blown diffuser can thus be made shorter than a conventional one. However, this generally requires that the blowing jet have a velocity considerably greater than the general tunnel stream at the location of the blowing slot. In a transonic tunnel, where the flow velocity at the entrance to the diffuser is near the speed of sound, a supersonic blowing jet would probably be required. This would introduce a great deal of noise into the tunnel circuit, which is undesirable for a number of reasons.
With regard to the overall problem of sustaining laminar boundary-layer flow, the so-called Griffith airfoils were first developed in the 1930's. The intent of these airfoils is to provide laminar boundary-layer flow over its entire surface in a manner to also achieve very low drag. The thought is that laminar flow could be maintained if the airfoil could be designed to have a favorable pressure gradient (accelerating flow) everywhere on its surface. A practically realizable airfoil shape with positive thickness, however, cannot have accelerating flow everywhere, and decelerating flow (a pressure rise) must be tolerated on the surface. The basic idea behind the concept is to shape the airfoil so as to concentrate the pressure rise in a very short region on the aft part on one or both surfaces. To prevent the boundary-layer separation that this sudden pressure rise would otherwise cause, a single boundary-layer suction slot is placed at the location of the pressure rise, where enough of the boundary-layer is removed so that the remaining flow can negotiate the pressure rise and attach itself to the surface aft of the slot. Two such airfoils are shown in FIGS. 1A and 2A, with FIGS. 1B and 2B showing the corresponding pressure distribution curves, these being taken from the work entitled "Incompressable Aerodynamics", D Thwarites, published by Dover Publications, Inc., New York 1958. Wind tunnel tests of the so-called Griffith airfoils demonstrated that the desired pressure distribution could be achieved in practice and that the minimum suction flow rate required to keep the flow attached agreed reasonably well with a theoretical criterion proposed by G. I. Taylor. However, the goal of full cord laminar flow was not realized because the concavity of the surface aft of the slot caused instability and transition to turbulence. Thus, although attached flow is achieved, the low drag levels originally predicted were not obtained.
Later, the same concept was applied to a body of revolution by F. Goldschmied, this being presented in "Integrated Hull Design, Boundary-layer Control and Propulsion of Submerged Bodies: Wind tunnel Verification", AIAA Paper 82-1204, AIAA/SAE/ASME 18th Joint Propulsion Conference, June, 1982. Wind tunnel tests of the Goldschmied body again showed that the type of pressure distribution required by the concept could be achieved in practice, and that Taylor's criterion again predicted the minimum suction rate reasonably well. FIG. 3A illustrates the Goldschmied configuration, with FIG. 3B presenting a pressure distribution curve for the same.
A feature added by Goldschmied was that a Ringloeb cusp would be helpful in achieving steady flow in that slot. In one of his tests, Goldschmied deliberately tripped the boundary-layer to turbulent near the nose of the body and found that the concept could be made to work with a turbulent boundary-layer upstream of the slot. The thicker turbulent boundary-layer merely required a higher suction flow rate, as predicated by Taylor's criterion. This particular configuration is shown In FIG. 4 of this patent application. Reference is also made to the work, "Separation Control by Trapped Vorticies" of F. O. Rinegloeb, published in the work "Boundary Layer and Flow Control, G. V. Lockmann, editor, Pergamon Press, 1984.
In an article "Tests in the N.P.L. Electric Tank of a 4:1 Axi-symmetrical Diffuser Having a Discontinuity in the Wall Velocity" by F. Cheers, and W. G. Rayner, published 1950 by "London: His Majesty's Stationary Office", there is disclosed a 4:1 axi-symmetrical diffuser which has a single velocity discontinuity in the surface. Such configuration is shown in both FIGS. 1 and 2 of this paper, where there is an upstream section of smaller diameter which expands in a relatively short axially length to a greater diameter in the downstream direction.
Also, there is a report entitled "An Improved Design Method and Experimental Performance of Two Dimensional Curved Wall Diffusers" by Tah-Teh Yang, W. G. Hudson and Ali M. El-Nashar of Clemsen University. This is a NASA report designated CR 121024, having a report date of Nov. 20, 1972. This report describes three different two dimensional diffusers, a so-called Griffith Diffuser shown in FIG. 12, a "dump diffuser" shown in FIG. 16, and a "cusp diffuser" shown in FIG. 17. These were constructed so that the two side walls were essentially planar, and there was boundary layer suction along these two sidewalls. The upper and lower walls have an upstream smaller depth dimension, and these diverge upwardly and downwardly in a relatively short axial length to a downstream location where is a substantially greater depth dimension. Also, there are upper and lower section slots provided in the transition region, and attention is called particularly to FIG. 12 where the "Griffith" diffuser is shown.
A search of the patent literature has disclosed a number of patent which are noted below. These are as follows:
U.S. Pat. No. 2,709,917 (Bruynes) shows a transonic wind tunnel where the transonic velocity in the test section is accomplished as follows. There is a fan which directs the air into a converging passageway upstream of the test section, where there is a tangential slot through which air flows along the boundary layer in a downstream direction at the upstream end of the test section. The wall 22 in the test section has a diameter greater than that of the air at the downstream end of the converging inlet. Air at the downstream end of the test section is drawn out as suction air 26 in a circumferential slot. Also, further suction air is drawn out of the air stream at a circumferential location 50 downstream of the circumferential slot 26.
U.S. Pat. No. 2,729,974 (Lee et al) shows a transonic wind tunnel where there is a converging inlet at the upstream end of the test section, and just downstream of this converging inlet, there is a tangential slot through which air is injected in a downstream area into the boundary layer of the air stream. In the test section itself, the circumferential wall 18 is formed as a porous layer, and there is suction boundary layer to draw the air into a circumferential chamber. Further, there is a pump 30 which draws the air out of the chamber 20 to inject it into the tangential slot to flow into the air stream.
U.S. Pat. No. 2,948,148 (D'Anferville et al) discloses a supersonic wind tunnel which can be operated at variable Mach numbers. The patent states that the common way of changing the Mach number in a supersonic wind tunnel is to change the cross-sectional area at an upstream location where the air is going from sub-sonic to supersonic. In this patent, the same effect is accomplished by injecting air into the air stream at the necked down location so as to aerodynamically restrict the flow at that area.
U.S. Pat. No. 3,000,401 (Ringles) discloses a boundary layer control device in the form of a cusp formed at an intermediate location. Upstream of the cusp, there is an aerodynamic surface which is contoured so that the air flowing from location A to C is accelerating slightly or at least is not substantially reduced at any point so that there is no boundary layer separation. This aerodynamic surface ends in a sharp circumferential edge facing in a downstream direction and away from the airflow, and there is, in cross-sectional configuration, a recess formed in a generally circular curve which extends from the sharp edge in a curve to join to a downstream aerodynamic surface. The effect is that within the recess, the air rotates about a fluid center at F and the airflow over the airfoil contour from A to C remains laminar. When this device is used in an airfoil, it is stated that there can be greater angles of attack, without the air separating over the aerodynamic surface from location A to C. Also, in column 4, beginning at line 26, the patent states that "The cusp effect described here can also be applied wherever a flow is to be led from a lower to a higher pressure or vice versa, as for example, to shorten the diffuser of a wind tunnel, or, in ducts and pipes as illustrated at 14 in FIG. 3 to prevent separation of the flow at turns and bends."
U.S. Pat. No. 3,049,005 (Frenzl) shows a wind tunnel where compressed air is taken from a tank 4 and fed through an ejector 3 which draws air from a testing chamber 1. The air from the ejector pump 3 along with the air drawn from the chamber 1 has a recirculating path 8, 9 and 10 back to convergent-divergent discharge nozzle 11. Downstream of the ejector pump 3 there is a second ejector pump 16 positioned downstream of the first ejector pump 3.
U.S. Pat. No. 3,981,144 (Milling et al) shows a dual stage supersonic diffuser which is designed to produce an effect similar to the change in the flow cross-section of a variable throat area diffuser. The diffusing element of the first stage is aligned perpendicularly to the second stage. The diffuser vanes of the first stage are inclined at an angle to the direction of flow so that the flow field between the first stage and the second stage changes from the conditions before diffuser starting to the conditions after diffuser starting.
U.S. Pat. No. 3,998,393 (Petty) shows a supersonic diffuser for use with a gas dynamic laser having a multi-channel shock duct diffuser with outer duct walls and inner duct walls forming an inner duct channel and two outer duct channels. In FIG. 8 and following, there are shown various separation point stabilizers, with FIG. 14 showing a modified version where there is a separation point at 51 which points in a downstream direction, with this being undercut as at 53 to form a recess having at least partially a curved configuration.
U.S. Pat. No. 4,515,524 (Fisher) shows a hydraulic turbine where there is a draft tube having a first diverging section to receive the water pumped from the turbine, and a further downstream section having a greater angle of divergence so that separated flow could occur. There are suction openings in the second section which draw off water and direct these back to an upstream nozzle location in the first diverging section, so as to alleviate separation In the second section.